The invention relates to after-burner devices for bypass turbojets, and more particularly it relates to improving the ignition performance of such devices.
Bypass turbojets for military airplanes are commonly fitted with an after-burner device. The after-burner device comprises an after-burner channel that receives from the turbojet both a “hot” central primary flow and a “cold” peripheral secondary flow, and that is connected at its outlet to a nozzle. The primary and secondary flows are obtained by splitting the total flow entering the turbojet into two flows. The primary flow passes through the high pressure compressor, the combustion chamber, and the high pressure and low pressure turbines of the turbojet, and reaches the after-burner device downstream from the low pressure turbine. The secondary flow flows at the periphery of the turbojet and it is used in particular for cooling certain members. The after-burner device further comprises means for injecting fuel into the vicinity of flame-holder members and an ignitor member, generally a spark plug, situated in an after-burner ignition zone. During after-burning, additional fuel is injected and is burnt by the oxygen contained in the two flows. This results in an increase in thrust.
It must be possible for the pilot to be able to ignite the after-burner device during any stage of flight and very quickly. Unfortunately, under certain flying conditions where the pressure generating the flow entering into the turbojet is low, and where consequently the pressure in the after-burner device is low, ignition performance can be insufficient. This is not compatible with the operational requirements of the airplane.